This disclosure relates to gas turbine engines and particularly to internally cooled turbine vanes.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
As is well known, the aircraft engine industry is experiencing a significant effort to improve the gas turbine engine's performance while simultaneously decreasing its weight. The ultimate goal has been to attain the optimum thrust-to-weight ratio. One of the primary areas of focus to achieve this goal is the “hot section” of the engine since it is well known that engine's thrust/weight ratio is significantly improved by increasing the temperature of the turbine gases. However, turbine gas temperature is limited by the metal temperature constraints of the engine's components. Significant effort has been made to achieve higher turbine operating temperatures by incorporating technological advances in the internal cooling of the turbine blades.
Various cooling passage configurations have been used to cool turbine vanes, but there may be some inadequacies in some applications. To this end, a double wall cooling configuration has been used to improve turbine vane cooling. In a double wall blade configuration, thin skin core cavity passages extend radially and are provided in a thickness direction between the core cooling passages and each of the pressure and suction side exterior airfoil surfaces. Double wall cooling has been used as a technology to improve the cooling effectiveness of a turbine blades, vanes, blade out air seals, combustor panels, or any other hot section component.
Turbine vanes need cooling on both the airfoil and the platforms. Conventional investment casting processes limit cooling designs to “pullable” cores, separating the airfoil and platform cooling circuits. Typical vane cooling design configurations will have cast cooling circuits on the airfoils and machined cooling holes on platforms. Given relatively flat combustor temperatures, significant flow is required to cool vane platforms and mateface gaps. Interstage gaps between the stationary and rotating turbine stages require a balance of sealing and purge flow to prevent hot gas from entraining into these areas and damaging the unprotected backsides of these components as well as non gaspath components with lower temperature capability. Introducing cooling air into the gaspath as a coolant imparts a cycle penalty reducing fuel efficiency of the engine, therefore making it desirable to utilize as little cooling air a possible to protect turbine components.